The present invention relates generally to gas turbine engines, and, more specifically, to cooling therein.
In a gas turbine engine air is pressurized in a compressor and mixed with fuel in a combustor for generating hot combustion gases which flow downstream through turbines. A high pressure turbine (HPT) first receives the hot gases from the combustor for extracting energy therefrom for powering the compressor through a corresponding drive shaft therebetween. A low pressure turbine (LPT) follows the HPT and extracts additional energy from the combustion gases for typically powering a fan in a turbofan gas turbine aircraft engine application. The LPT in industrial and marine engine applications may otherwise be used for driving an external drive shaft.
In order to ensure a suitably long life for the engine, the various combustor and turbine components directly subject to the hot combustion gases are typically cooled during operation by bleeding a portion of compressor air for cooling thereof. Cooling in a gas turbine engine is esoteric and quite sophisticated and uses various forms of holes for channeling the cooling air as required for the disparate turbine components requiring cooling.
For example, film cooling holes are common in gas turbine engines and have various configurations typically being inclined through the wall being cooled, with an inlet for receiving the compressor cooling air and an outlet for discharging that air to generate a thin film or boundary layer of air which thermally insulates the wall from the hot combustion gases flowing thereover. The air channeled through the holes provides internal convection cooling of the component wall, with the film generated on the outer surface of the wall effecting a thermally insulating barrier.
Furthermore, it is common to introduce a thermal barrier coating (TBC) in selected areas of various engine components for providing additional thermal insulation for limiting the temperature of the underlying metal or substrate wall being protected.
Combustor liners and turbine components are formed of various superalloy metals having sustained strength at the high temperatures of operation experienced in a gas turbine engine. Efficiency of operation of the engine is maximized by maximizing the temperature of the combustion gases generated therein, but that temperature is limited by the ability to suitably cool all engine components along the flowpath of the hot combustion gases.
Cooling design in a gas turbine engine is made more complex since the engine is operated over various power levels including steady-state and transient operation in which temperature correspondingly changes. And, the combustion gases vary in temperature according to position, and correspond with varying heat transfer loads into the differently shaped engine components.
For example, the typical combustor in a turbofan gas turbine engine includes an annular dome having a plurality of circumferentially spaced apart carburetors in which fuel injectors are mounted in corresponding air swirlers. A fuel and air mixture is discharged from each of the carburetors and ignited for generating the hot combustion gases which correspondingly effect a temperature variation around the circumference of the combustor. Each carburetor therefore generates a locally hot streak of combustion gases which is reduced in circumferential variation at the outlet end of the combustor, yet such circumferential variation in combustion gas temperature nevertheless exists.
The various engine components must therefore be designed for suitable cooling for the hottest expected combustion gas temperature even though those components will normally operate with lower temperature combustion gas. The overall efficiency of the gas turbine engine is therefore limited by the need to provide a nominal amount of cooling air for each component notwithstanding the variation in need for cooling air.
Accordingly, it is desired to provide a flowpath wall in a gas turbine engine having improved cooling features therein.
A flowpath wall is provided for bounding hot combustion gases in a gas turbine engine. The wall includes opposite outer and inner surfaces and a plurality of cooling holes extending therebetween. A thermal barrier coating is bonded to the outer surface and covers blind the holes thereat.